1. Industrial Field of the Invention
The present invention relates to an improvement of a turbine blade in a gas turbine and, more particularly, to a cooling structure of the turbine blade.
2. Description of the Relative Art
By burning fuel with an oxidizing agent of high-pressure air which has been compressed by a compressor, a gas turbine serves to drive a turbine by high-temperature high-pressure gas thus produced, in order to convert the generated heat into energy such as electricity. As a method for improving the performance of a gas turbine, working gas has been changed to have higher temperature and higher pressure. When the temperature of the working gas is elevated, it is necessary to cool a turbine blade and maintain its temperature not to exceed a practical temperature of material of the turbine blade. An example of a conventional cooling structure of a turbine blade is disclosed in ASME, 84-GT-114, Cascade Heat Transfer Tests of The Air Cooled W501D First Stage Vane (1984), FIG. 2.
In this cooling structure of the turbine blade, the blade is of a double structure, i.e., the blade body has a hollow-structured body provided with an inner constituent member (hereinafter referred to as the core plug) therewithin. A large number of apertures are bored through the core plug so that compressed air extracted from a compressor is discharged from these apertures (hereinafter referred to as the impingement holes) against the inner surface of the blade body, thus performing impingement cooling by strong impingement air jets. The air which has cooled the turbine blade from the inside is discharged from the suction and pressure sides or the trailing edge of the blade into main working gas. The number of the impingement holes at each location is appropriately chosen in accordance with fluid heat transfer conditions of the main working gas, thereby allowing the whole blade to have a substantially uniform temperature. The exterior surface of the blade in the vicinity of the leading edge is exposed to the gas of high temperature, which has a particularly high heat transfer rate there. This leading edge portion has a curvature which is unfavorably large for cooling, and accordingly, the cooled area of the inner surface of this portion is relatively small in comparison with the heated area of the outer surface of the same. Therefore, a great number of impingement holes are located inside of the leading edge portion so as to cool it with a large amount of cooling air. This tendency has been especially strengthened in response to the recent elevation of the gas temperature.
Another example of a conventional cooling structure of a turbine blade in a high-temperature gas turbine is disclosed in ASME, 85-GT-120, Development of a Design Model for Airfoil Leading Edge Film Cooling (1985), FIG. 1. In this cooling structure, the blade is of a double structure equivalent to the above-described conventional example, where impingement cooling is conducted by discharging cooling air from impingement holes of a core plug within the blade, and also, film cooling is performed by releasing part of the cooling air into main working gas from a large number of apertures (hereinafter referred to as the film cooling holes) formed at a portion in the vicinity of a leading edge portion of the blade.